1. Field of the Invention
The present invention relates generally to gas turbine engine, and more specifically to turbine rotor blade with integrated cooling and sealing for use in a gas turbine engine.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine, such as a large frame heavy duty industrial gas turbine (IGT) engine, includes a turbine with one or more rows of stator vanes and rotor blades that react with a hot gas stream from a combustor to produce mechanical work. The stator vanes guide the hot gas stream into the adjacent and downstream row of rotor blades. The first stage vanes and blades are exposed to the highest gas stream temperatures and therefore require the most amount of cooling.
The efficiency of the engine can be increased by using a higher turbine inlet temperature. However, increasing the temperature requires better cooling of the airfoils or improved materials that can withstand these higher temperatures. Turbine airfoils (vanes and blades) are cooled using a combination of convection and impingement cooling within the airfoils and film cooling on the external airfoil surfaces.
In the prior art, near wall cooling utilized in an airfoil mid-chord section is constructed with radial flow channels plus resupply holes in conjunction with film discharge cooling holes. As a result of this cooling design, spanwise and chordwise cooling flow control due to the airfoil external hot gas temperature and pressure variation is difficult to achieve. In addition, single radial channel flow is not the best method of utilizing cooling air resulting in a low convective cooling effectiveness. The dimension for the airfoil external wall has to fulfill the casting requirement. An increase in the conductive path will reduce the thermal efficiency for the blade mid-chord section cooling. FIG. 1 shows a cut-away view of a prior art turbine blade with near wall cooling. FIG. 2 shows a cross sectional view of the blade with two radial flow cooling channels in the pressure side and suction side walls. The blade mid-chord section is cooled using multiple single pass radial flow channels 11 each having an oval cross sectional shape. Film cooling holes 12 connect the radial channels 11 to the external surfaces of the airfoil. Cooling air from one or more cooling air supply channels 13 formed within the airfoil through resupply holes 14 and into the radial channels 11. In the design of FIGS. 1 and 2, the cooling through flow velocity as well as the internal heat transfer coefficient is comparatively reduced. Subsequently, refresh holes along the internal wall of the radial flow channel is used to replenish the cooling flow.